Hypersonic velocity leading edge



Dec. 17, 1963 c. E. CONN, JR.. ETAL 3,114,524

HYPERSONIC VELOCITY LEADING EDGE 5 Sheets-Sheet 1 Filed Aug. 21, 1961 R S mm T R m m w w M 8' F KS AN A CH V. B

ATTORNEY Dec. 17, 1963 c. E. CONN, JR.. ETAL 3,114,524

HYPERSONIC VELOCITY LEADING EDGE 5 Sheets-Sheet 2 v Filed Aug. 21, 1961 FIG. 2

INVENTORS CHARLES E. CONN JR.

BY HANS F. MUELLER M FEM ATTORNEY Dec. 17, 1963 c. E. CONN, JR, ETAL 3,114,524

Dec. 17, 1963 c. E. CONN, JR. ETAL 3,114,524

HYPERSONIC VELOCITY LEADING EDGE 5 Sheets-Sheet 4 Filed Aug. 21, 1961 IO N INVENTORS CHARLES E. CONN JR.

BY HANS F. MUELLER M FEM,

ATTORNEY Dec. 17, 1963 c. E. CONN, JR.. ETAL HYPERSONIC VELOCITY LEADINGEDGE 5 SheetsSheet 5 Filed Aug. 21, 1961 SUPPORTING STRUCTURE (WING) LEADING EDGE STRUCTURE PIN JOINT CONNECTION- wmakmmmzwh wo nEDm INVENTORS CHARLES E. CONN JR. BY HANS F. MUELLER M FTDMV ATTORNEY United States Patent Oilice 3,ll4,52 l Patented Dee. l7, 1963 3,114,524- HYPERSQNTC Vl'ELilCllY LEADlNG EDGE Charles E. Quinn, J12, Manhattan Beach, and Hans 1F. Mueller, Playa Del Rey, Qalitl, assignors to North American Aviation, Inc.

Filed Aug. 21, 1961, Ser. No. 132,953 6 Claims. (Ell. 244-l17) This invention relates to structures for withstanding thermal shocks and high temperatures and in particular it relates to an aerospace vehicular structure including a nonablative leading edge portion capable of withstanding hypersonic velocities that may occur either during cruise or during entry into a planetary atmosphere.

The design of aerospace vehicles for hypersonic cruise velocities or for reentry into a gaseous atmosphere such as surrounds the earth and some of the planets poses severe structural problems due to the very high temperature and thermal stresses imposed on the vehicle structure. The availability of materials that can successfully withstand such extreme thermal environmental conditions is severely restricted due to the oxidation, erosion, melting and general reduction in physical properties which most materials undergo when exposed to such extreme temperatures and high degree of thermal shock. In general, materials, such as graphite, ceramics, ceramic oxides, and refractories, possess properties enabling them to withstand such severe environmental conditions. Most of these materials have low ductility. This is particularly true With respect to ceramic oxides which have excellent temperature and oxidation resistant properties compatible with a hypersonic leading edge environment. Eecause of this property of brittleness, the design of a structure capable of optimally utilizing such materials presents difficult problems since it is necessary to insure that the structure has suflicient freedom to prevent failure from thermal stresses as well as the ability to Withstand high dynamic air loads and severe vibration.

The present invention provides an uncooled, nonablative type leading edge structure largely employing radiation from solid surfaces and conduction r'or temperature and heat control. A basic concept of the present invention is, firstly, the isolation of the higher temperature leading edge structure from the supporting structure. Thus, the nonablative leading edge structure is made separate from the supporting structure and attached thereto in a manner LO effectively thermally isolate the leading edge structure from the supporting member and greatly reduce heat conduction into the latter member. Secondly, the present arrangement includes the making of the leading edge structure from a material which is many times thicker than the skin of the supporting structure or has a larger coetficient of thermal conductivity or both. As a consequence, a substantial quantity of heat will be conducted from the stagnation region of the leadins edge structure to the downstream end of the structure where it is effectively dissipated by radiation due to the increased surface temperature in that area.

While a unitary one-piece leading edge structure of the above type may sutlice for certain temperature levels, for higher performance requirements the leading edge structure itself is also preferably comprised of material segments zoned in a chordwise direction in accordance with their temperature and oxidation limits and having sufiicient relative freedom to prevent failure from thermal stresses. Thus, more specifically, the present invention provides a bimaterial leading edge or nose portion comprised of a graphite base member having forwardmost ceramic nose portions pinned thereto with freedom for relative expansion between the individual ceramic nose segments and between the nose segments and the graphite base member. The invention further includes a heat short for equalizing the temperatures of the extremities of the base member and serving as a thermal radiation shield for the adjacent structure supporting the leading edge.

Accordingly, it is a primary object of this invention to provide a leading edge or nose structure capable of Withstanding the thermal shock, high temperatures and environmental conditions occurring during hypersonic tlight of an aerospace vehicle through a planetary atmosphere.

Another object of this invention is the provision of a leading edge structure capable of withstanding hypersonic velocity and orbital reentry temperatures which is effectively thermally isolated from its associated supporting structure.

Another object is the provision of a combination ceramic-graphite leading edge structure wherein the foremost portion of the structure is comprised of segmented ceramic members joined to a graphite base member Without the introduction of clamping stresses.

Still another object of this invention is the provision of a combination ceramic-graphite leading edge reentry structure wherein the foremost portion of the structure is comprised of segmented ceramic members joined to a graphite base member in a manner whereby freedom for relative expansion is provided between the individual ceramic and graphite members.

A further object is the provision of a bimaterial leading edge reentry structure wherein the leading edge or nose portion is comprised of ceramic segments having overlapping joints that allow expansion but prevent air-flow through such joints.

Yet another object is the provision of a reentry-type leading edge structure having a rear wall portion acting as a heat short and as a radiation shield to protect the adjacent lower temperature-resistant materials associated therewith.

It is also a still further object to provide a leading edge reentry structure that is pin-connected to a supporting metallic structure whereby the supporting structure is substantially thermally insulated from the leading edge structure.

These and other objects and advantages of this invention will become apparent to those skilled in the art after reading the present specification and the accompanying drawings forming a part thereof, in which:

FIG. 1 is a perspective view of a portion of an aerospace vehicle having an airfoil incorporating the preferred embodiment of the leading edge structure of the present invention.

FIG. 2 is a top plan view of a leading edge structural segment in accordance with the preferred concept of the present invention.

FIG. 3 is a chordwise sectional view through the structural segment of FIG. 2 showing such segment in assembled relationship with the supporting airfoil structure.

FIG. 4 is a perspective exploded view of the segmented bimaterial leading edge structure of the present invention.

FIG. 5 is a graphical illustration of the time-temperature relationship experienced during a plasma jet tunnel test of a leading edge segment made in accordance with the present invention.

FIG. 6 is a graphical representation of the chordwise temperature distribution existing along a unitary pinattached leading edge structure.

A leading edge, nose cone or other forwardly projecting portion of a hypersonic aerospace vehicle is normally subjected to a temperature far greater than the temperature imposed upon the remainder of such a vehicle upon its passage through a relatively dense atmosphere. Such conditions exist to the maximum extent during reentry of an orbit-M vehicle into the earths atmosphere. it has heretofore been normal design procedure to attempt to provide for such conditions by incorporating in the structure a leading edge or frontal portion of a heat resistant material which is capable of sustaining the extremely high temperatures generated. The leading edge portion is normally terminated and fixedly attached to the primary structure at a sufiicicnt distance downstream from the leading edge so that the design temperature at that location would not be so high as to impair the strength of the primary structure material.

Such a leading edge, nose cone or other forwardly exposed portion has, as the result of its hypersonic velocity through the earths atmospheric gases, a temperature gradient existing therethrough of gradually decreasing temperatures with increasing distance rearwardly away from the stagnation point temperature which exists at its most forward point. At the extremely high stagnation point or stagnation line (locus of stagnation points) temperatures encountered during an atmospheric reentry by an orbital vehicle (as high as 300i) F. to 4%0 F. or more dependent on the vehicles trajectory, velocity and geometry, such as leading edge radius and wing sweep) the chordwise temperature gradient causes a gradient in the spanwise expansion which introduces compressive and tensile stresses within a continuous skin. Thus, an expansion joint in the skin would serve to localize the thermally induced stresses and reduce the heat conducted across the joint to a small fraction of that conducted through a continuous skin. Such a joint should be provided where the temperature differences reach an upper limit which is determined by the thermal and elastic properties of the leading edge material. The present invention provides joints acting as heat barriers which also allow sufiicient freedom for expansion to localize the build-up of thermal stresses. This minimizes the effects of thermal shock and prevents failure of the leach ing edge structure from thermal stresses. For less rigorous reentry conditions where the temperatures to be encountered may not be too severe, the leading edge structure may be of homogeneous material and/or of unitary construction having only one thermal insulating joint between the supporting structure and the leading edge unit at the point of attachment. For more severe reentry conditions and higher temperatures the use of additional joints at the forwardmost nose portion of the leading edge structure in accordance with the preferred embodiment of FIGS. 2, 3, and 4 is advantageous. Furthermore, the segmented nose arrangement of this preferred embodiment of the invention permits zoning or the use of different materials having temperature, erosion, strength and oxidation-resistance properties compatible with the environmental conditions to be encountered at a particular area or zone on the leading edge structure. It also permits easy replacement of individual damaged segments.

A number of materials, including graphite, molybdenum, columbium, chromium, tungsten, tantalum, rhodium, thorium oxide (thoria) and beryllium-oxide (beryllia), exhibit properties which permit their use in structures of the above described type. Some of the physical properties of the importance for such an application are the melting point, strength, modulus of elasticity, coefficient of expansion, thermal conductivity and specific weight. The ratio of thermal conductivity to specific weight is also of particular interest, since it is desirable to conduct as much of the heat as possible downstream and away from the immediate vicinity of the stagnation point by selecting materials of high thermal conductivity, large thickness or a combination of both. This aids in effectively dissipating the heat by radiation due to the increased surface temperatures at the rearward extremities of the leading edge structure and it also reduces the stagnation point temperatures and the chordwise temperature gradient.

Of the listed materials, graphite has the largest ratio of thermal conductivity to specific weight and is, in general, a preferred material for a temperature-resistant leading edge structure. Graphite, however, is susceptible to oxidation and erosion and must be protected by a suitable protective coating. Carbide, nitride and silcide ceramic coatings are among those available for this purpose. Graphite, also, has the further disadvantage of possessing a very low tensile strength. Thoria and beryllia, however, are highly refractory ceramic materials that retain usable mechanical strength at elevated temperatures and are unexcelled for stability at high temperatures in oxidizing atmospheres. in addition to a high degree of chemical inertness, they possess low sublimation vapor pressures at high temperatures and thus may be used under vacuum conditions. Thus, the use of a pin joint for attachment of the leading edge structure of the present invention allows for expansion of such structure and significantly reduces the transmission of heat and propagation of thermal stresses into the supporting metallic structure. The pin attachment arrangement also permits taking advantage of the high thermal conductivity and other factors, such as mass, which are conducive to heat flow through the structure by insulating the leading edge rearmost extremities from the supporting structure. This raises the temperature of such extremities and increases the thermal radiation therefrom.

For more severe conditions of temperature and thermal shock, the further use of a segmented nose arrangement on the leading edge structure itself permits expansion of such nose segments relative to the base structure of the leading edge unit and further localizes the thermal stresses and prevents their transmission and propagation into the base structure of the leading edge unit. The. latter (arrangement is particularly advantageous since it permits the combination of various materials in a manner whereby desirable properties of such materials may be utilized to form an optimum composite leading edge structure. Such a composite structure, for example, thus may consist of a graphite base structure in combination with nose portions of a efractory ceramic material.

Referring specifically to the drawings, wherein like reference characters I ave been used throughout the several views to designate like parts and referring at first to FlG. '1, reference numeral 1 generally designates an aerospace vehicle having an airfoil surface 2 for the purpose of permitting control of reentry through a plane tary atmosphere and precise landing of the vehicle on the surface of the planet. During such reentry the en tire vehicle is subjected to a rapid build-up of tempera ture due to aerodynamic heating with the most forwardly exposed portions, the nose and leading edges, being ex posed to the maximum temperatures, i.e., the local stagnation temperatures cheated by passage of the vehicle through the relatively dense atmosphere at hypersonic velocities. For a typical reentry trajectory, such stagnation temperatures may rapidly rise from an initial temperature of perhaps 603 1F. to 4200 F. or higher. Besides imposing severe temperature resistance requirements on the leading edge materials and structures, the rapid imposition of such extreme temperatures also induces severe thermal shock in such materials. T0 withstand such environmental conditions, a high temperature leading edge structure 3 is provided on the airfoil by means of a series of leading edge segments 4-.

FIGS. 2 and 3 present a top plan view and a sectional view, respectively, of a preferred embodiment of one of such leading edge segments. As shown therein, the segment comprises a base structure 5 which includes relatively thick chordwise extending upper and lower walls 6 and 7, respectively. The upper and lower Walls are joined at the for-ward portion of the base structure by front end wall 9 and at the rear of the structure by a rear end wall 8 which, in addition to helping to form an integral box-like structure, has a unique temperature control function to be further described hereinbelow.

Projecting forwardly from the upper portion of front wall 9 is a lip 11 which serves as a seat for the ceramic nose or leading edge segments 20. At the rearward extremities of the base member an arrangement is provided whereby the leading edge structure can be pin-connected to the associated supporting structure which normally would consist of the front spar of an airfoil such as a Wing or similar structure of the vehicle. As shown in FIG. 3, this arrangement includes a lug 12 extending rearwardly from the junction of top wall 6 and rear Wall 8. Lug 12 contains a spanwise extending aperture 13 and is divided by a series of slots 14 which permit the reception therein of complementary lugs or ears formed on the metallic supporting structure St A similar arrangement is utilized at the junction of lower wall 7 and rear wall 8 to form a lug 15 having a spanwise aperture 16 and slots 17 formed therein. Upon fitting of the leading edge segment to the supporting structure 39, a rigid conne tion is made by the insertion of pins 32 and 33 into apertures 13 and 16, respectively.

Since the longitudinal axis of the structure during reentry is angularly displaced from the airstream by an amount equal to the angle of attack, the lower surface of the leading edge structure is generally at a higher temperature than corresponding portions of the upper surface of the structure. To equalize the temperatures and reduce the thermal gradients in the upper Wall member, the rearward ends or extremities of the base structure have been joined by rear wall 8 which acts as a conductive path or heat short tending to reduce the maximum thermal gradient and equalize the temperatures between such extremities. In addition, this wall also serves as a shield against heat radiation protecting the front spar 3i of the metal supporting structure 39. The heat short serves to reduce thermal stresses in the leading edge structure by as much as thirty percent.

Attachment of the nose segments, which are generally extremely hard and brittle, is provided by a plurality of ceramic pins interconnecting the nose segments to the base member of the leading edge structure. The pins are angularly disposed relative to each other in pin holes which are sufi iciently oversized to allow for slight relative expansion between the pins and the bores. In this fashion, no screw or bolt-type attachments need be used with the concomitant clamping type stresses which would be introduced thereby. As shown in FIG. 3, pin 23 extends in a chordwise direction and is contained in matching bores Zll and 18 in the nose segment and base member, respectively. A canted pin 24 is similarly contained in canted bores 22 and 19 formed in the leading edge segment and the lip ll of the base structure, respectively. This canted pin arrangement is approximately 45 to the chordwise extending pin 23. Bore 19 terminates at its upper end in a plug hole 25 normal to the surface of Wall 6. Hole 25, in turn, is fitted with a graphite plug as which secures canted pin 24 in the structure.

Each base structure segment has a plurality of leading edge nose segments 2%, as shown in FIG. 2. The abutting edges of the leading edge segments 2d are formed to provide a loosely fitted scarfed tongue-in-groove type joint 27 which allows for unrestricted spanwise expan sion between the nose segments but prevents airflow through the joints.

in assembling the structure, the horizontal chordwise ceramic pin is first installed in the ceramic leading edge segment Zil and this assembly is then installed on the graphite base structure by insertion of the exposed portion of the pin into bore 13. The canted ceramic pin is then inserted through plug hole 25 into matching canted holes 19 and 22. The insertion bore is then plugged with a press fit graphite retainer plug 25. In the exploded view of FIG. 4-, pins 23 are shown for purposes of illustration inserted in the graphite base structure hole 18.

FIG. 6 graphically presents the advantages accruing from the use of a leading edge structure 35 formed of a relatively thick material of high thermal conductivity which is pin-connected to the supporting structure 35. While a unitary leading edge structure is illustrated therein, the following discussion and advantages are equally applicable to the preferred segmented arrangement of the present invention. As a result of the proper selection of a high conductivity material of adequate thickness, a substantial quantity of heat will be conducted from the stagnation region 37 to the rearward extremities 33 where it will be more effectively dissipated by radiation due to the increased surface temperature in that area. Curve A demonstrates the chordwise equilibrium temperature distribution for such a structure as compared with that for a thin-skin integral leading edge (broken line curve B). It is clear that such an arrangement simul aneously reduces the temperature in the stagnation region from T to T and also the chordwise temperature gradient from AT to AT Reduction of the temperature gradient results both from the use of a relatively thick leading edge member having a high thermal conductivity as well as from the thermal isolation of the leading edge from the supporting structure by the pin attachment. The former arrangement tends to flatten the thermal curve of temperature vs. chordwise distance while the pin connection raises the temperature of the rear extremities of the leading edge while maintaining the supporting structure at a safe, low temperature. Reduction of the chordwise temperature gradient is most desirable in that it reduces the number of span-wise expansion joints required in the preferred form of construction.

While the present invention has been described in detail relative to an airfoil leading edge, the applicability of the concept and structure presented herein is equally adaptable to any edge portion of a structure of the stated type which permits the use of a connecting pin structure. It will be clear to those skilled in the art that, while the segmented leading edge nose portion has been disclosed as applied to a symmetrical leading edge, it is equally applicable to an asymmetrically configured leading edge such as is disclosed in a copending application. Similarly, it will be clear to the artisan that the selection of a graphite material for the base structure and a beryllium-oxide leading edge segment or nose portion in the particular embodiment described herein in no way limits the concept set forth in this specification. The present invention provides a means whereby any suitable materials may be joined in a manner to provide a composite structure having a high degree of resistance to extremely high temperatures and the thermal stresses imposed thereby.

Tests have been successfully conducted in a plasma jet of a leading edge structure made in accordance with the preferred embodiment of the present invention, as shown in FIGS. 2, 3 and 4. Such tests prove the ability of this device to withstand temperatures in excess of 4000 F. for sustained periods of time Without damage, as shown by test curve C of time versus temperature in FIG. 5. This curve closely approximates the expected time-temperature profile for a typical reentry mission. The test structure consisted of three 1" long beryllia segments having a symmetrical nose radius of /s". These segments were pin-connected to a siliconized graphite base structure to form an assembly 4% long and 3" wide. The model was held in the plasma at zero sweep (normal to the plasma jet) with a 10 angle of attack. Surface temperatures and temperature distributions along the model were obtained by means of optical pyrometers, and color motion picture cameras were used to record the nature and time of any changes in the structure. These tests proved the feasibility of the concept and structure of the present invention for withstanding great extremes of temperature and thermal shock.

Thus, the present invention provides a hypersonic cruise or reentry type leading edge structure advantageously utilizing joints and pin-connected structures to control heat flow and localize the build-up of thermal stresses.- It permits the use of a base material such as graphite for a primary material while protecting it from extreme stagnation temperatures and temperature gradients tothereby reduce oxidation, erosion, and vaporization of the graphite. It permits the use of different materials in combination whereby such materials may be advantageously zoned to utilize the optimum physical properties of each such material to produce the most effective combination. Thermal stresses are also further controlled by the provision of means for unrestrained expansion of the primary heat sustaining members, i.e., the nose segments, and by a heat short member forming a thermal flow path across the rearward ends of the graphite base member to equalize the temperature at the inner ends of the upper and lower surfaces, reduce the maximum.

temperature gradient and serve as a heat-radiation shield protecting the metallic structure to which the leading edge assembly is attached.

While a particular embodiment of this invention has been illustrated and described herein, it will be apparent that various changes and modifications may be made in the construction and arrangements of the various parts without departing from the spirit and scope of this invention in its broader aspects, or as defined in the following claims.

We claim:

1. A structural member adapted to form a forwardmost portion of a hypersonic velocity vehicle capable of withstanding the thermal stresses and stagnation temperatures generated during hypersonic flight through a gaseous atmosphere comprising a hollow non-metallic base member having apertured lugs at its upper and lower rearmost edges which are adapted to be pin-connected to and thereby substantially thermally isolated from a supporting metallic structure; a plurality of non-metallic nose segments mounted on the forward end of said base member; a plurality of pins connecting each said nose segment to said base member with sufficient clearance between said nose segments and said base member whereby said nose segments have three-dimensional freedom of expansion, said pins being angularly disposed relative to one another whereby said nose segments are substantially fixed attached to said base member.

2. A structural member, as set forth in claim 1, wherein said plurality of pins comprise the sole means securing :said nose segments to said base member and further comprise a chordwise extending pin and a canted pin angularly disposed to said chordwise extending pin.

3. A structural member, as set forth in claim 2, wherein said chordwise extending pin and said canted pin are wholly embedded within and secure together said base member and said nose segments.

4. A structural member, as set forth in claim 1, wherein said base member and said nose segments are formed of refractory and ceramic materials, respectively, and wherein the joints between said ceramic nose segments overlap in a spanwise direction to prevent airflow through the joints while allowing unrestrained expansion of the nose segments and wherein said supporting metallic structure, said base member and said nose segments form a :smooth aerodynamically faired leading edge surface.

5. A structural member, as set forth in claim 4, wherein said hollow refractory base member is tapered toward its leading edge terminating in a forwardly extending lip portion at its upper leading edge surface, said nose segments having a rearwardly extending lower lip portion complementary to said base member forwardly extending upper lip.

6. A structural member, as set forth in claim 5, wherein said hollow refractory base member has a lower surface greater in width than the upper surface and further having a refractory thermal barrier wall sloping upwardly and forwardly to join the rearmost ends of the base member lower and upper surfaces, said barrier wall being adjacent to thermally protecting the front edge of the metallic supporting structure by substantially short circuiting the flow of heat to the supporting structure and shielding such structure from radiation from other portions of said base member while tending to substantially equalize the temperatures of the rearmost ends of said base member upper and lower surfaces.

References Cited in the file of this patent UNITED STATES PATENTS 2,446,766 Hosbein et al. Aug. 10, 1948 2,473,728 Rutledge June 21, 1949 2,535,883 Williamson Dec. 26, 1950 3,028,128 Friedrich Apr. 3, 1962 

1. A STRUCTURAL MEMBER ADAPTED TO FORM A FORWARDMOST PORTION OF A HYPERSONIC VELOCITY VEHICLE CAPABLE OF WITHSTANDING THE THERMAL STRESSES AND STAGNATION TEMPERATURES GENERATED DURING HYPERSONIC FLIGHT THROUGH A GASEOUS ATMOSPHERE COMPRISING A HOLLOW NON-METALLIC BASE MEMBER HAVING APERTURED LUGS AT ITS UPPER AND LOWER REARMOST EDGES WHICH ARE ADAPTED TO BE PIN-CONNECTED TO AND THEREBY SUBSTANTIALLY THERMALLY ISOLATED FROM A SUPPORTING METALLIC STRUCTURE; A PLURALITY OF NON-METALLIC 